1. Technical Field
This invention discloses generally a method of positioning a solar wing of a satellite relative to the sun, and more particularly, a method of positioning a solar wing of a satellite including compensating for the ephemeris motion of the sun by means of a sun sensor on the satellite's body.
2. Discussion
A geosynchronous earth orbit, as is known in the art, is the orbit about the earth in which a satellite or spacecraft will remain fixed above a specific location on the earth. This orbit is at a distance of approximately 22,400 miles above the earth. In this orbit, a beam, such as a communications beam, from the satellite can be maintained over a desirable area on the earth, such as a particular country, thus establishing an area which will receive the beam. To remain in a geosynchronous orbit it is necessary that the satellite be in an orbit substantially within the equatorial plane of the earth at the desirable distance, and that the satellite's attitude be oriented perpendicular to this plane. Any deviation or disturbance which causes the satellite to direct its antenna away from a boresight location on the earth tends to effect the coverage area of the beam, and thus, produces undesirable results. Many different forces are in effect on the satellite which tend to alter the satellite's antenna pointing direction.
As a first order method for countering the effects of the different forces acting on the satellite, it is known to stabilize the satellite's attitude by providing an angular bias momentum which resists changes in the satellite's orientation due to external forces transverse to the bias momentum axis. Satellites using this technique are generally referred to as "momentum bias" satellites. Angular momentum bias is usually provided by a number of momentum or reaction wheels which spin at least part of the satellite. The bias axis set by the spin of the momentum wheels is generally perpendicular to the direction of the orbit of the satellite. Although the bias momentum resists changes in the satellite's orientation in directions transverse to the bias momentum axis, it is still necessary to provide control for correcting variations in the satellite's orientation along the bias axis. Different methods of controlling the satellite's attitude, such as feedback loops, are known in the art.
For most bias momentum satellites, the satellite payload, i.e., the part of the satellite carrying at least the antenna, is oriented differently than the momentum wheel. It is therefore necessary to provide means for correcting the orientation of the payload with respect to the orientation of the momentum attitude. Typically, the satellite's payload is defined in three axes referred to as the yaw, roll and pitch axes. If the satellite is in a geosynchronous orbit, the yaw axis is generally directed from the satellite to the center of the earth, the pitch axis is generally directed normal to the plane of the orbit of the satellite and the roll axis is generally perpendicular to the yaw and pitch axes, in a direction of travel of the satellite as is well known in the art.
Satellites of the type discussed above generally include solar wings in order to generate the power necessary to operate the different electrical systems on the satellite. For maximum efficiency, it is well known that the solar panels on the solar wings need to be perpendicular to the direction of the incoming rays of light from the sun. However, at a geosynchronous altitude the apparent azimuthal position of the sun, i.e., the angle about the satellite pitch axis between the sun and the earth's center, as measured at the same time of day, changes seasonally up to about .+-.4.5.degree.. If this change in the position of the sun relative to the solar wings of the satellite is not corrected for, a significant reduction in efficiency of power collection by the solar wings occurs over time due to the fact that the light rays will be hitting the solar wings at a substantial angle from normal.
The prior art solar wing steering mechanisms on satellites generally include only means to track the sun at a constant rate. Consequently, as discussed above, this has allowed for certain tracking errors due to the sun's ephemeris motion in that the normal tracking rate does not consider this motion. In these types of systems, it has been generally required that sensing and adjusting of the wing position relative to the sun be done from a ground location to correct for this motion. This process provides a costly interference with the satellite which need not occur.
What is needed then is a system for automatically controlling the desired position of the solar wings of a satellite by means of a sensor on the satellite, without any intervention from a ground station. It is accordingly an object of the present invention to provide such a system.